METHOD OF AIRPLANE PERFORMANCE ESTIMATION AND PREDICTION

Abstract: A method of predicting performance characteristics of an aircraft including modeling thrust-minus-drag of an aircraft with at least one mathematical model using aircraft specific data and input parameters which define the performance characteristics. Coefficients of the at least one mathematical model are learned over at least one flight. The performance characteristics are predicted as a function of the learned coefficients representative of a combined and non-independent thrust-minus-drag relationship.

Description: METHOD OF AIRPLANE PERFORMANCE ESTIMATION AND PREDICTION FIELD OF THE INVENTION The present invention relates to flight management systems. In particular, the present invention pertains to a method of estimating, and subsequently predicting, aircraft performance. BACKGROUND OF THE INVENTION In the past, a "rule of thumb" method was utilized in predicting aircraft performance characteristics. With this technique, the pilot, based on previous experience, makes a reasonable estimate of the aircraft performance characteristics. Other past aircraft performance prediction systems include a custom performance computer which utilized "average parameters" for a predetermined model of an aircraft. This average parameter method is useful for only one aircraft type, and must be repeated for each new type of aircraft. These methods and systems have other inaccuracies associated therewith. For example, the rule of thumb method is very inaccurate since it is difficult to compensate for temperature of the air, weight of the aircraft, etc. Also, thi method usually adds to the pilot workload at a time when the pilot is already very busy. The custom performance computer method, in addition to the average parameter disadvantage, does not account for manufacturing tolerances, a degradation of parameters resulting from age and usage, nor for different pilot techniques. U.S. Patent No. 5,070,458 to Gilmore, et al., entitled "Method of Analyzing and Predicting both Airplane and Engine Performance Characteristics," issued December 3, 1991, describes a method which overcomes many of these disadvantages. This method makes performance predictions for an individual aircraft and engine using parameters which are learned from "flight to flight." The data, which is used by this method for predicting performance characteristics, is initialized with reasonable values of specific performance parameters. This method includes adjustment or updating of the specific performance parameters resulting from each flight of the aircraft. Thus, the performance characteristics of a given aircraft are learned from flight to flight for use in future flights of the given aircraft. The learned parameter technique thereby adjusts to changes in the aircraft due to aging, is tailored to a specific aircraft, and accounts for manufacturing tolerances. However, in Gilmore, et al., the modeling utilized and the computation of terms thereof are inefficient. In addition, Gilmore, et al. uses a model which separates thrust and drag terms of the model when making predictions and is somewhat limited in the number of outputs produced. For example, Gilmore does not provide for prediction of long range cruise speed or optimum altitude, which are important for achieving the best possible fuel efficiency. In view of prior estimation and prediction systems, a need is apparent for a prediction system which provides an enhanced set of outputs for use by the pilot. In addition, the outputs, which may include outputs previously computed by prior systems, should be provided with greater speed and/or improved accuracy. SUMMARY OF THE INVENTION The present invention provides an estimation and prediction method which provides an enhanced set of outputs which are provided with greater speed and/or accuracy. An estimating method for use in predicting performance characteristics of an aircraft in accordance with the present invention includes modeling the aircraft with at least one series expansion using aircraft specific data, such as wing reference area, and input parameters, such as current weight or altitude, which define the performance characteristics. Coefficients of the at least one series expansion are learned over at least one flight so that the learned coefficients can be utilized for predicting performance characteristics of subsequent flights. A thrust and drag relationship may be modeled wit a first series expansion and climb fuelflow of the aircraft may be modeled with a second series expansion. A first and second filter are utilized for estimating the coefficients of the first and second series expansions, respectively. In another method of the present invention, a method of predicting the performance characteristics of an aircraft includes modeling the aircraft with at least one mathematical model using aircraft specific data and input parameters which define the performance characteristics. Coefficients of the mathematical model are learned during at least one flight and cruise performance characteristics are predicted as a function of the learned coefficients of the mathematical model. Long range cruise speed, maximum cruise speed and optimum altitude at both such speeds are predicted. These performanc characteristics are computed using the mathematical model and a given set of input parameters, e.g. long range cruise speed at a given weight and altitude. These computations are optimized by differentiating functions derived from the mathematical model with respect to each independent input parameter. In another method of the present invention, parameters such as total fuel burned for a given flightplan are computed with minimum iterations utilizing closed-form integration. The ability to differentiate the mathematical model is necessary to such integration. An additional embodiment of the present invention includes a method of predicting performance characteristics of an aircraft including modeling thrust-minus- drag of an aircraft with at least one mathematical model using aircraft specific data and input parameters which define the performance characteristics. Coefficients of the at least one mathematical model are learned. Performance characteristics are predicted as a function of the learned coefficients representative of a combined and non-independent thrust-minus-drag relationship. The certain performance characteristics are a function only of the non-independent or combination thrust-minus-drag relationship, and not of thrust and drag independently. These predictions are of considerably greater accuracy than those depending on explicit knowledge of thrust and drag separately. BRIEF DESCRIPTION OF THE DRAWINGS Figure 1 shows a general block diagram of the method of airplane performance estimation and prediction in accordance with the present invention. Figure 2 shows a block diagram of the signal conditioning block of the learning portion of the method of aircraft performance estimation and prediction method of Figure 1. Figure 3 shows a block diagram of the performance estimation block of the learning portion of the aircraft performance estimation and prediction method of Figur 1. Figure 4 is a chart showing an aircraft flight. Figures 5A-5C are graphs showing variations in key parameters as a function o altitude. Figure 6 is a graph showing short trip limited altitude. Figure 7, including Figures 7A-7C, show method blocks for determining speed accordance with the present invention when altitude is fixed. Figure 8, including Figures 8A-8C, show method blocks for determining maximum and optimum altitude for manual, long range cruise, and maximum speed in accordance with the present invention. Figure 9 is a graph showing a specific range as a function of speed. Figure 10 is a graph utilized for determining maximum speed. Figures 11 A and 1 IB are graphs showing determination of optimum and maximum altitude of manual speed. Figure 12 is a contour graph showing thrust-minus-drag as a function of altitud and true airspeed. Figure 13 shows a specific range surface defined for determining long range cruise speed. Figure 14 shows a thrust-minus-drag surface defined for determining long rang cruise speed. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT The method of aircraft performance estimation and prediction 10 shall be described in general with reference to Figure 1. The aircraft performance estimation a prediction method 10 includes two major portions, a learning portion 12 and a predicti portion 14. The learning portion 12 includes modeling a thrust-minus-drag relationshi of an aircraft with a Taylor series expansion utilizing signal input parameters 26, such as current aircraft weight and altitude, and aircraft specific data, such as wing reference area, stored in non-volatile memory 32. Additionally, fuel flow at the climbing rate is modeled with another Taylor series expansion utilizing signal input parameters 26 and aircraft specific data. Performance estimation 16 receives conditioned signal inputs 28, and preferably utilizing Kalman filter techniques, estimates the coefficients of the Taylo series expansions of both the modeled thrust-minus-drag relationship and climb fuel flo over the course of multiple flights. The learning portion 12 yields performance estimate outputs 30 including fuel flow estimate vector x ff , thrust-minus-drag estimate vector x, and covariance matrices P and P ff , of which P provides data for determining how accurate the system will predict aircraft performance characteristics utilizing figure of merit block 24 of prediction portion 14. The learned coefficients of the Taylor series expansions for the thrust-minus-drag relationship and fuel flow along with their respective covariance matrices are retained in non-volatile memory 32 and are therefore available during subsequent flights to refine the existing estimates utilizing learning portion 12 and to predict aircraft performance characteristics with climb predictions block 18, cruise predictions block 20 and descent predictions block 22. The prediction portion 14 utilizes Newton's-method-related techniques to find certain points of interest on curves or surfaces representative of functions derived from the thrust-minus-drag relationship and fuelflow to be maximized or otherwise localized given particular constraints. Aircraft performance characteristics such as long range cruise speed, maximum speed, maximum and optimum altitudes and combinations thereof are examples of characteristics which are computed in this manner. Newton's- method-related techniques rely heavily on the ability to differentiate various functions of interest. This ability in turn depends ultimately on the ability to differentiate the thrust- minus-drag function and fuelflow which are expressed as Taylor series expansions whos coefficients are provided directly by performance estimation 16 and whose derivatives are readily available. With the learned coefficients of the Taylor series expansions and respective covariance matrices retained in non-volatile memory 32, the data can be downloaded an utilized for initializing or "seeding" a new example aircraft of similar type. For example, a first aircraft can be utilized in the learning process, while the data learned can be used to seed the learning process of a second similar type aircraft or used for predictions of the second aircraft characteristics. In addition, with the coefficients and covariance matrices in non-volatile memory, the pilot is allowed to save the data if a box of the system is pulled for any reason. It can then be reloaded when desired mitigating the nee to completely restart the estimation process. With reference to Figures 1-14, the present invention shall be described in furthe detail. The non-volatile memory 32 includes an aircraft data file which is used for performance predictions once sufficient learning has taken place and also for continuing the learning process. Sufficient learning in the learning portion 12 has taken place to begin the prediction portion 14 usually after about 3-4 flights. The aircraft data file is divided into two parts. The first file part contains data that is independent of learning, but still aircraft specific. This aircraft specific data includes the following: - aircraft type - aircraft tail number - ceiling altitude (feet), h ceil - default climb constant calibrated air speed (KCAS) - default climb Mach - default descent KCAS - default descent Mach - takeoff fuel allowance (lbs) - landing fuel allowance (lbs) - basic operating flight weight (BOW) (lbs) - wing platform area (ft**2), S - maximum operating airspeed, Vmo (KCAS) - maximum operating mach number, Mmo - maximum lift coefficient, C, - critical mach number, M cr The second file part of the aircraft data file contains the learned coefficients of both the Taylor series expansions and their respective covariance matrices. The learned data of the second file part of the aircraft data file includes: - 11 coefficients for the thrust -minus-drag Kalman filter - 5 coefficients for the rated-fuelflow Kalman filter - the upper triangle of the eleven-by-eleven covariance matrix including the diagona elements describing the thrust-minus-drag filter (a total of 66 values) - the upper triangle of the five-by-five covariance matrix including the diagonal elements describing the rated-fuelflow filter (a total of 15 values). As indicated above, the aircraft- is modeled by use of Taylor series expansions of thrust-minus-drag and rated fuelflow about a given reference point. For the remainder the present description, given a Taylor series expansion of the function z = f{x,y) of th form: ^ ( ^^ ) 0 ,^ +^ ,^, ( ^-^ + U = , 0 ( - ) = f 0 +f x, --x+f yo Ay then " f o -f x -f y > etc * wι -- b- 5 referred to as coefficients of the series, - Δx, Δy (or x and y with the "-x 0 " understood), etc. will be referred to as parameters of the series, and - f x Δx, etc. are terms of the series. The terms "coefficients" and "parameters" are also used in the context of a general mathematical model. The Taylor series expansion for the thrust portion T of the thrust-minus-drag relationship is expressed as: where Mis Mach number, δ is pressure ratio, W f is corrected fuelflow defined further below, h p is pressure altitude, and \ 0 indicates the preceding term is to be evaluated at the midpoints M 0 , W f and h p<> which are also further defined below. The drag portion D of the thrust-minus-drag relationship is expressed as: D = 1481<5M 2 S C D = C_M 2 C Ό where C* = 1481S S = wing reference area C D = coefficient of drag The coefficient of drag C D is modeled as: C D = C Di (l +T) where C a = coefficient of drag-incompressible flow r = compressibility correction as described below. The incompressible drag coefficient -, is modeled as: where = coefficient of lift C D ,C D ,,&k are coefficients to be learned. Since lift L can be written as CβM 2 C L and is equal to weight Win flight, drag D can be expressed as: W 2 D-=(C l SM 2 C Do +WC Di +-^ T k)(l + T) The compressibility correction term T is expressed as: T = 0 when M ≤ M cr T = \5(M-M cr ) 2 tan[( - c _) ( 3 +.l)A] when M>M cr where M cr - critical mach number λ = coefficient to be learned Combining the expressions for thrust and drag, the thrust-minus-drag relationship T- D is modeled as follows: T-D=δ[x_+ x 2 AM + x^W fc + x 4 Ah p + x 5 Δ 2 + x 6 ΔMAW fc + x η AW ] where x.= l,d x [ 0 , x 2 = - τiδ) i , x,= %τ/δ) '., x 4 = jiδ) X.._ aΛ δW f ch p *" * 8 = D„ ' 9 ~ ^D_ ' * .(. ~ " ■ > X ** — /I Taken collectively, the coefficients x, , x 2 ,etc. can be expressed in vector form as x = [x, , x, , ] . The performance estimation portion 16 of the present invention develops estimates of these true coefficients; the estimates denoted by x ϊ ,x 2 ,et r ' or by the vector x. The 11-term Taylor series expansion allows more degrees of freedom thar. ;. the prior art, resulting in the modeling of significant higher order effects which yield better results. Also, the altitude h p becomes an intrinsic part of the modeling in the present invention, unlike other methods of estimation and prediction which utilize ar "after the fact" curve fit of altitude with other parameters. The reference point about which the Taylor series is expanded is chosen for best fit over typical operating conditions. In the preferred embodiment, it is defined to be M B = 0.60 mach, h„ = 30,000 feet, and W, = 7,4001bs/hr, thus: AM = M - 0.60; Ah = h p - 30,000; AW fc = W fc - 1,400 The rated fuelflow, also modeled by a Taylor series expansion, is expressed as: where T DJSA = T '- T 1SA ; T m being standard temperature, or may be expressed as: W fr = x ff _ + x ff2 AM + x ffi Ah + x ff T D + x ffi AhT Dm where X ff. - W fr ' x X ff* - ~ ^rJrL { I" ' x X f - * W * ■ c '° f *-- ~ cT ch p w The rated fuelflow W fr is modeled by the Taylor series expansion about the same reference point as thrust-minus-drag is modeled. Taken collectively, the coefficients x ff .... x ff can be expressed in vector form as x ff - \x ff x ff ] r . The performance estimation portion 16 of the present invention develops estimates of the coefficients; the estimates denoted by x^ .... x^* or by the vector x ff . With specific reference to Figures 2 and 3, the learning portion 12 of the present invention shall be described in further detail. Input signal conditioning is provided by integrator 34, corrected fuelflow block 36, atmospheric model 38 and two steady-state Kalman filters, altitude rate filter 40 and acceleration rate filter 42. Input signal conditioning block 15, Figure 2, receives the signal parameter inputs 26 including fuelflow W f , mach number M_ temperature T, pressure altitude h p _ pitch θ, roll φ, sensed longitudinal acceleration a LON , sensed normal acceleration a NOR and true air spee V τ . The inputs W f _ M_V τ ,andh p are also inputs to performance estimation block 16. Fuelflow W f is applied to integrator 34 along with initial aircraft weight W 0 , determined from specific aircraft data including BOW and other pilot input as known to one skilled in the art, to provide an estimated weight W as fuel is utilized. Corrected fuelflow block 36 for output of corrected fuelflow W f is provided with inputs of fuelflow W f , mach M, and temperature T. Corrected fuelflow W fc is defined by: where£ is the pressure ratio and θ is temperature ratio. Pressure ration and temperature ratio θ are provided to corrected fuelflow block 36 from atmospheric mod block 38 which has the inputs of temperature T and pressure altitude h p . The standard atmospheric relationships which determine these inputs to corrected fuelflow block 36 are well known and understood by those skilled in the art. By means of integrator 34, corrected fuelflow block 36 and atmospheric model 38, conditioned inputs 28 including estimated gross weight W _ corrected fuelflow W fc pressure ratio δ and the deviation from standard temperature 7^ are provided to performance estimation block 16. The altitude rate filter 40, preferably a steady state Kalman filter, provides the estimated output of altitude rate h p . The general equation for a Kalman filter is written as x k+ = Kz + Mx κ with z being the measurement vector, K being the gain matrix, M being a matrix derived from the gain and transition matrices, x λ . +1 being a (2,1) new state vector and x k being the previous state vector. The only input to the altitude rate filter 40 is the pressure altitude h p ; z vector for the filter is of dimension (1,1). The K and M matricies are given by: K = The (2,1) new state vector x t+l , consists of estimates of altitude and altitude rate. The desired altitude rate output h p is the second element of this new state vector. The acceleration filter 42, also preferably a steady-state Kalman filter and generaliz in the same manner as altitude rate filter 40, provides a smooth estimated output rate o airspeed V τ . However, simply computing the time rate of change of V τ from the airspe input V τ is insufficient. Thus, the filter also uses inertial accelerations. The acceleratio in the stability reference frame is given by: cι x = cos a(a WN - sin θ) + sin a(-a N0R + cos-^cosø) where the angle of attack a is computed from: , ( h . a = θ- γ where γ = sin " — ) with θ being pitch angle. The measurement vector z of this filter is: =[κ * x ] The K and M matricies for this filter are given by: .0847 .5915 .9153 .3242 K = M = .0054 .1797 -.0054 .8122 The desired output of rate of airspeed V τ is the second element of the new state vector for this filter. As described above, the conditioned inputs 28 and several signal inputs 26 are provided to performance estimation block 16. They include gross weight estimate W, corrrected fuelflow W fc , pressure ratio δ, pressure altitude h p , deviation from standard temperature T DISA , fuelflow W f , altitude rate h p , airspeed rate V τ , and true air speed V τ . Performance estimation block 16, Figure 3, utilizes two Kalman filters to learn the coefficients of the Taylor series expansions modeling thrust-minus-drag and rated fuelflow as previously described. Performance Kalman filter 48 is utilized to learn the coefficients of the thrust-minus-drag relationship and fuelflow rating Kalman filter 52 is utilized to learn the coefficients of the Taylor series expansion modeling rated fuelflow. The performance Kalman filter 48 is an extended Kalman filter necessary for estimating coefficients of a non-linear function. Kalman filtering techniques are known to those skilled in the art. This conventional approach to filtering involves propagation of a state estimate vector and a covariance matrix from stage to stage. It should be readily apparent to one skilled in the art that an appropriate filtering technique may be used in this performance estimation portion of the invention to provide the process of learning coefficients of a mathematical model. Such filtering could also be utilized where filters are used in signal conditioning block 15. Likewise, it is recognized that neural networks could be used in lieu of the series expansion modeling technique. There are five recursive equations that make up a Kalman filter used in the present invention for filters 48,52. Preferably, they are run on a 10 second cycle as long as qualifying measurements are available. The state vector is expressed as x . The covariance matrix P describes the degree of uncertainty in x . The process is assumed to be subject to zero-mean, Guassian white noise des c ribed by a Q matrix which is diagonal The measurement vector is z and its associated co . αnance matrix is given by R. Since, in the cases of the performance Kalman filter 48 and the fuelflow rating Kalman filter 52, essentially steady-state coefficients are being estimated, the respective state transition matrices Φ become identity I, and the standard Kalman filter equations reduce as follows: - propagation equations: x " „+ι = Φx k = x k P ~ ' +i = ΦP Φ τ +0 - = R - k +0 - gain computation: K k = P ~ - H τ (HP ~ k H τ + R)-' = P kH HP ~ k H τ +R since R is scalar in these cases. - update equations: P k = (I - K k H)P ~ k x k = x ~ k + K k (z k - Hx ' k ) = (/ - K k H)x k ~ + K k z k In accordance with the general description of Kalman filters above, the T-D state estimate vector x is of dimension (11,1) including the coefficients of the T-D Taylor series expansion as discussed previously. The state vector x is given by: Therefore, the update vector H is of dimension (1,11) and consists of the parameters of the T-D series. The update vector His computed as shown by block 46, Figure 3, as a function of Mach number M, pressure ratio δ , weight W _ corrected fuelflow W f and altitude h p . The update vector H is provided to performance Kalman filter 48. The first seven elements of update vector H are given by the thrust update vector H τ and the last four are given in the drag update vector H D . The thrust update vector H τ is as follows: H r = φ, AM, AW fc , Ah, (AM) 2 , AMAW fc , (AW fc ) 2 ] The drag update vector H D is Mach number dependent. The incompressible H D o H D is: where C, is (1481)(S) . The compressible H D or H Dc is only computed when the Mach number Mis above the critical Mach number M c and is given by: H D = [θ, 0, 0, ( h x,)2T ξ ≥ 1 sin (2 )λ where and where the summation run is from i of 8 to 10. Finally, when M < M cr then H D = H D and when M > M -r t/7e/7 H D = H (l + r) - H^ Because of the greatly differing ranges of the variables in update vector H, it is normalized by a matrix N. This normalized update vector H' is then used in the Kalman filter equations. This vector H' is defined as: H'= HN where the normalization matrix, N, is a diagonal matrix of dimension (11,11). The diagonal elements are defined as: N = [3000, 2000, 0.1, 0.005, 2000, 0.1, 0.00001, 0.005, 0.005, 0.03, 25] As shown by block 44, the measurement vector, z, of dimension (1,1) and representative of thrust-minus-drag T-D is computed from the gross weight estimate W and the inputs of true airspeed V τ , altitude rate h p and rate of airspeed V τ as follows: where g is the gravitational constant of 32.174 ft/sec 2 , f j is a knots to ft/sec conversion factor of 1.6878065, and f2 is a ft/sec to knots conversion factor of 0.5924850. With the associated error covariance R (1, 1) of measurement vector z set to a predetermined vector and the diagonal elements in Q (11,11) set to a predetermined vector, x', the normalized form of x, can be computed. Not-normalized x is then computed from the Kalman filter output of x and N as follows: x = N ' In order to take a measurement of the quantity thrust-minus-drag T-D, an estimate o current acceleration must be available. In a cruise condition of the aircraft, current acceleration is zero and a measurement is not taken unless this condition is ensured. However, in climb or descent of aircraft acceleration is not zero. In prior art systems, acceleration was estimated from a planned speed schedule; however, this schedule could be violated in actual flight. The present invention bases its acceleration estimate on -** currently estimated values of rate of change of true air speed V τ and altitude rate h p utilizing inertial acceleration components. Also, measurements are processed only when estimated acceleration is with acceptable limits. The following conditions determine when the thrust-minus-drag equation is valid and consequently when to run the Kalman filter for the learning process: - The airplane must be 5,000 ft above the takeoff runway. - The bank angle must be less than 5 degrees. - The rate of change in pitch must be less than .1 degree per second. - The Mach number must be greater than 0.40. After has been learned, the coefficients thereof can be used in the prediction portion 14 of performance estimation and prediction method 10 in subsequent aircraft flights. The Kalman filtering approach propagates the covariance matrix rather than information which is the inverse of the covariance matrix. This leads to a much more computationally efficient form of filter. With use of normalization and double-precision arithmetic more maintainable software is achieved without compromise of accuracy. Although the individual terms of the mathematical model can be separated as to which apply to thrust and which apply to drag, the estimate of the combined relationship of thrust-minus-drag (T-D) is more accurate than individual estimates of thrust and drag as they are not independent. Greater accuracy can be achieved in prediction of characteristics by expressing most of the computations required for prediction in a form that does not depend on such separation. With regard to fuelflow estimation, the fuelflow rating Kalman filter 52 is run when the following filter conditions are satisfied. The core filter equations for the filter 52 have already been discussed with regard to performance Kalman filter 48. The conditions include: - The aircraft must be more than 1,500 ft above the takeoff runway. - The aircraft vertical speed must be greater than a minimum vertical speed V πaD defined as (units of ft/min): - The aircraft must be controlling to airspeed while climbing and at the same time be 1,000 ft below any preselected level off altitudes. - The fuelflow must not have changed more than 400 lbs/hour since the last 10 second sample of the filter. If it has, a 30 second timer is started before learning is again enabled. The fuelflow state estimate vector x ff is of dimension (5,1) and is defined in accordance with the partial derivatives of the Taylor series expansion modeling fuelflow. The state vector x ff is given the subscript ff to distinguish it from the thrust-minus-dra state estimate vector and is given by: x ff = \ x ff. - - x ff. \ The update vector H ff is computed as a function of Mach number M, pressure altitude h p and deviation from standard temperature T D1SA as shown by H ff matrix computation block 50. Update vector H ff is of dimension (1,5) and is defined by: H ff = \ 1, AM, Ah p , T Dm , T asA h p ] Because of the greatly differing ranges of the variables in H ff , it is normalized by matrix N^ resulting in normalized H' ff which is used in the Kalman filter techniques an is given as follows: H' J ~ = H ff N ff where N^* is a diagonal matrix of dimension (5,5); the diagonal elements thereof given by: N = [l000, 300, 0.05, 10, 0.005] The measurement vector z ff (1,1) is the current unfiltered and uncorrected fuelflow W f . With the associated error covariance R ff of z ff set to a predetermined vector and the diagonal elements in O ff (5,5) set to predetermined values, x' ff , the normalized vector can be computed. Not-normalized x ff is then computed from the Kalman filter output of x' ff and N^- as follows: The vector x ff can then be utilized by predicting portion 14 of aircraft performance estimation and prediction method 10. With rated climb fuelflow W f being modeled as a function of Mach number M, altitude h p and deviation from standard temperature at altitude T DISA , significantly improved estimates of fuel consumed during climb are achieved. In addition, the results of this estimation are employed to predict characteristics such as maximum cruise speed. The state estimate vectors x ff and x and the covariance matrices P ff and P are par of the aircraft data file and stored in non-volatile memory 32. Each time a transition from airborne true to false is detected, an aircraft flight counter is incremented and the file updated with the latest values for such matrices. Only if a new aircraft data file is loaded is the data replaced; the learned data from one aircraft being capable of usage for seeding other aircraft. The initial values of x ff for the first flight in a new aircraft type are set to: x o = [4000, 1000, -0.15, -25, O.Ol] The initial values of x 0 for the first flight on a new aircraft type are set to: x 0 = [12000, 0, 2, 0, 6000, 1, 0, 0.02, -0.01, 0.1, 150], and the P D and P ff matrices are set to identity matrices. Data from a first application of the method on a new aircraft type creates a data file to be used as a starting file or seeding file on other aircraft of the same type. The x ff matrix, the x 0 matrix, and the diagonal elements of P 0 and P ffo will remain unchanged; however, all non-diagonal elements in the P 0 and P ff matrices are set to 0. With a computer containing the present invention including an aircraft data file that i sufficiently learned, various predictions can now be accomplished via the predicting portion 14 of the performance estimation and prediction method 10. Such various predictions are made with regard to climb, cruise, and descent and as a function of modeled thrust-minus-drag and/or modeled rated-fuelflow. An aircraft flight is divided into three phases: climb phase, cruise phase and descent phase, Figure 4. The climb phase occurs from the origin of the flight to top of climb (TOC). The cruise phase extends from TOC to top of descent (TOD). Finally, the descent phase begins at TOD and continues through destination. Predictions for these phases of the flight are made via the method blocks: climb prediction 18, cruise predictions 20 and descent predictions 22. Climb predictions 18 provide the following information: time enroute, t, fuel remaining F r , altitude h p , airspeed V τ , and groundspeed V c at each waypoint in a flightplan. Also, the top of climb predictions including the distance-to x τoc and fuel remaing at top of climb F τoc are determined. The prediction methods involve extensive use of first order Taylor Series expansions. The quantities climb angle γ, rated fuelflow W fr , and true airspeed V τ , are all assumed to vary linearly with altitude as shown in Figures 5A-5C. They are written as: dV V- = V T +~ t ^ l (h -h B ) dh where γ T- D Y / - Y / „ ai * r r — -rr ~ - and * r a_i.r_ = W(\ + ACF) where ACF is a standard acceleration factor known to those skilled in the art, the aircraft gross weight, V G is groundspeed, and subscript 0 means the quantity evaluated at a reference altitude. The climb angle γ is referenced to earth fixed coordinates and is therefore corrected for the effect of winds. The rate of change of γ ajr with respect to altitude h p is written: By assuming that the wind increases with altitude, the rate of change of climb angle γ with altitude h p is also corrected for the effects of wind by: dW f „ The rate of change of rated fuelflow with respect to altitude — — is readily available dh p from the rated fuelflow W fr equation previously discussed. Similarly, the rate of change of thrust-minus-drag with respect to altitude — is available from the thrust- dh p minus-drag T-D equation previously discussed. The rate of change of true airspeed with dV altitude — - can be found by assuming either a constant calibrated airspeed or a constan dh p mach number and either a constant lapse rate or a constant temperature. Distance x, time t, and fuelburn, F, are now written as integrals with respect to altitude h p in forms which allow the solutions to be found using standard integral tables. The integrals are expressed as: x = \-dh p r W, F = i fr h.. v τ Tyf , air The climb integration is typically subdivided into three steps. These steps are: a lower step where it is assumed that the aircraft flies at a constant calibrated airspeed, a middle step where the aircraft flies at a constant Mach number but is below the tropopause, and finally an upper step where the aircraft is flying at a constant Mach number but is above the tropopause. The step altitudes are determined by a specific aircraft's climb KCAS/Mach defined in the aircraft data file or entered by the pilot. For dV dV the — - = 0 and for the two other steps, — - can be determined assuming a standard dh p dh p lapse rate. The cruise altitude chosen, either an optimum altitude or a pilot entered altitude, may be set low enough to preclude one or more of these steps. The starting altitude for climb predictions is always the aircraft altitude. Thus, as the aircraft climbs, the number of climb steps are reduced. Once in cruise, no climb predictions are made. The reference altitude for each step, where the various quantities are evaluated, is always chosen to be the step's mid-altitude. Distance x as a function of altitude h p which is used for TOC placement, becomes: where h e is the altitude at the end of the step, h b is the altitude at the beginning of the step, and By rewriting this equation, altitude h p as a function of distance x, can be found for use in making altitude predictions at w /aayyppcoints, or in this case at the end of a step The solutions to time t and fuelburn F are subdivided into two parts, one when V τ is constant and second when V- is changing with altitude h p . For c and Εor dV τ /dh p ≠ 0 where, With fuelburn F, time t, and distance determined for each step, values at top of climb such as x τoc and F τoc are simply the sum of all the individual steps. Cruise predictions 20 provide estimates of time enroute t TOD and fuel remaining at TOD F TOD and distance to TOD x TOD . The description provided below for cruise predictions of total cruise fuelburn F and time t is described with respect to a single leg flight plan; however, it is equally applicable to a multiple leg flight plan in which case predictions are provided at each waypoint in cruise. In cruise, the altitude h p remains constant, but the aircraft's gross weight W might change significantly depending upon the length of the flight plan. Required fuelflow W f decreases correspondingly as weight W decreases. In addition, the groundspeed V a can be influenced significantly by the winds, and to a much lesser extent by temperature changes. Furthermore, when flying either at the maximum speed M ^ or at the long range cruise speed M LRC , the mach number M (or the calibrated airspeed) will change with weight W. Like the climb prediction method, it is assumed that fuelflow W f , airspeed V τ , and winds vary linearly. This allows for use of first order Taylor series expansions, which again produce integrals that can be solved using standard techniques. First, fuelflow W f is known as a function of weight Wϊvom the thrust-minus-drag T- D equation by setting T-D equal to zero. The rate of change of fuelflow with respect to dW f weight — — can then also be found using standard methods for computing derivatives. Thus, fuelflow W f can be written as: where W f = initial fuelflow at TOC and W 0 - initial gross weight at TOC from climb predictions block 18. Cruise time as a function of weight can be written as the following integral: t = J— dW W f This integral can be solved to yield time as a function of weight, but the useful form is weight or fuelburn as a function of time. Casting the equation in the form of fuelburn F as a function of time t, yields Distance traveled x as a function of time is given by the following integral: x = IV a dt where V G = V τ cos( l) - W H with λ = drift angle and W H is the average component of the wind along the ground track. V τ is related to the Mach number M, assuming that the Mach number M changes linearly with time as follows: dM t V τ = a\ where "a" is the speed of sound at the cruise altitude and for an average cruise leg temperature. An expression for the average drift angle is found by integrating over the crosswind, assuming again a linear change as expressed by: where W c is the component of wind normal to the ground track with W c = beginning W c and W c = ending W c . Time then becomes the solution of a second order equation: Mach number M changes with time depending upon both temperature and wind at the beginning and the end of a cruise leg. The method employed to determine is as follows. A first estimate of cruise time, assuming is zero, is used to find time t. dt This allows computation of fuelburn F, and predicted Weight W, at the end of the leg or as in our case the end of the cruise phase. Then, the Mach number M at the end of the leg is computed at the first estimate of W. can then be evaluated. If a constant Mach number cruise mode is chosen, as discussed later herein, this step is not necessary. With computed, the time t and fuel estimates are then refined. dt The effect of a changing wind W H can be significant because the aircraft will spend more time in the headwind than it will in the tailwind. The above estimates of time were determined using an average W H . When the wind changes significantly, a slowdown factor /_ can be estimated by: The original time estimate can now be multiplied by the slowdown factor /_ to give the final time and fuelburn values. Descent predictions 22 are performed in a similar manner to the climb predictions 18. For descent predictions block 22, which utilizes x, the inputs of a descent schedule, winds, and estimated cruise altitude h P„en , two basic modes can be utilized. The first mode is called a path descent method where the ground referenced descent angle γ d is constant. This method fixes the TOD placement regardless of winds predicted. The other method fixes the descent angle γ d and consequently moves TOD based on predicted winds. For short flight plans, the optimum altitude h p may be limited by how high the aircraft can go before a descent is necessary. The following rule described with referenc to Figure 6, is employed: the cruise phase should be at least 30 nautical miles, but no more than one quarter of the total flight plan distance —x lol . A corrected distance x Λsl , 4 can be constructed accounting for the origin and destination elevation and subtracting th cruise distance. The short trip limited altitude, h s , can then be estimated. It is assumed that the descent angle γd is constant and that the climb angle γ c decreases linearly with altitude. Using the standard Sine Rule with a small angle approximation and the first order Taylor series expansion for the climb angle γ c leads to the following second order equation for h s : \o d χ d, , rcY = ( - This equation can then be solved to yield h s which provides the short trip limited optimum cruise altitude. Along with the predictions of fuelburn, time, and other performance characteristics such as climb angle, predictions for speed and/or altitude in various speed modes of cruise predictions block 20 as shown in Figures 7A-7C and Figures 8A-8C are performed utilizing Newton's method related techniques. Newton's method related techniques allow for determining certain points of interest on curves that represent functions to be maximized or otherwise localized given particular constraints. This leads to the ability to compute speed in various speed modes as shown in Figure 7A-7C with altitude fixed. They include long range cruise speed, Figure 7 A; maximum speed, Figure 7B; and maximum endurance speed, Figure 7C. It also leads to the ability to compute altitude with speed fixed as in manual speed at optimal altitude, Figure 8A. In addition, in higher dimension cases, Newton's method related techniques are used to find certain points of interest on surfaces that represent functions to be maximized, such as where speed and altitude are variable. Such modes include long range cruise speed at optimal altitude, Figure 8B, and maximum speed at optimum altitude, Figure 8C. In the case of long range cruise, speed and altitude are determined for providing best fuel mileage, or in other words, the speed which maximizes specific range S R is determined; specific range being defined as: S D = -^~ W f wherein V G is ground speed and W f is fuelflow. With mach M being determined as a function of fuelflow W f while altitude and weight are fixed, and T-D is taken to be zero or some small positive constant, Newton's method related techniques shall be described for calculation of long range cruise speed M IRC , Figure 7A. To describe calculation of long range cruise speed, a curve, Figure 9, is shown wherein the x-axis is mach number M and the y-axis is specific range S Λ which is a function of mach number M and fuelflow W f . Maximum range cruise mach M mrc is the mach number for which S R (M) is a maximum. Long range cruise Mach M mc is the point above M mr- at which S Λ (M) =.99S Λ (M mrc ). M mrc is determined in accordance with the following steps with altitude p and weight f being given parameters: a) Choose an initial estimate, M, : where C, = 1481(S), Wis gross weight and δ is atmospheric pressure ratio. b) Compute the specific range S R and the derivative of specific range S^ with respect to mach at this point: The details of this computation are described further below. c) Choose a second initial estimate, M 2 : where M. = mach step size = .05 initially d) Compute S Rι and S R __ in the manner like explained below for S R and S R .. e) Estimate the second derivative S R f) Check if the second derivative is non-negative. If non-negative, the current "best fit" parabola has a minimum so no maximum will be found. This implies that the real S R function has an inflection point and the search region must be moved past the inflection point which is also away from the current minimum: where M 1A - mach "inflection adjustment" = .2 If the second derivative is negative, the current best-fit parabola has a maximum which is "close" to the real M mrc . This maximum is the point at which S Λ (M) = 0: M,Si - M 2 S Ri M. ->.„' ~ = ■ Ol _ o * Ensure that this new value for M mrc is not farther than M 1A from M, . If so, set M mκ = M, ± M 1A whichever limits the movement of M mrc without changing its direction. g) Reduce the mach step size M_ = M. 12. h) If the absolute value of M mrc - M, >.003, revise the initial estimate: M_ 1 = M mr r c c and repeat steps b)-g), otherwise go to step i). i) Compute S„ = S R (M mrc ) and S' = S R ' (M mrc ). j) Estimate the second derivative at this final maximum: Off - ** * •AC -A , k) Compute the long range cruise speed: The computation of specific range S R as for a given value of mach M, and of the first dS„ derivative, S R ' (M) = — —(M), step b) above, given specific values of pressure altitude dM h p , temperature deviation T D]SA , weight Wand current coefficients x are as follows: a) Compute AM = M- M 0 and Ah p = h p -h Po where M 0 =.β and h Po = ~ 0,000 ft b) Compute temperature ratio θ and pressure ratio δ from h p and T DISA by standard atmospheric relationships. c) Compute true airspeed V τ : V τ ~ ~ - a 0 ^Θ M where a 0 is the standard sea-level speed of sound = 661.4748 kts. r xτr ~ d) If M > M , perform steps e)-g) to compute T, — , and , otherwise set dδ dM „ dT dT _ . . , , . r = — = = 0 and continue with step h). dδ dM v e) Compute coefficient of lift C L : W r = - L C,δM 2 where C, =1481(S) f) Compute two interim terms, ς and ξ : ς=\5(M-M cr ) 2 ; ξ=(M -M 2 (C L '+Λ) g) Compute T, — , and SJ V dδ dM r = ςt ~ \ ~ (ξ X U ) h) Thrust-minus-drag T-D as a function of corrected fuelflow W fc is written as a quadratic in AW fc : T-D = aAW, - +bAW f +c=T where T n is residual climb thrust and is given by: where h rc is taken (in this case) to be \00fpm and the quadratic coefficients are computed as: a = δ x η b - <5(x 3 + x 6 AM) c=-T rc + δ(x { + x 2 AM + x 4 Ah p + x^ M 2 )-h D - where h D = x M 2 + x 9 W + x 10 —=rpr)Q + T) cδM- i) Find the solution to the quadratic: AW f = if\a\> ε, otherwise AW fc --clb f - 2a f where ε is a small positive constant. j) Find fuelflow for this mach number: W, = AW, +W, , where W, = 7400 Ibslhr. k) Find groundspeed V G and specific range S R : ,V G -< V T ~-W_ where W H = headwind -» ^ R W f 1) Compute the partial derivatives of fuelflow with respect to mach and to correcte fuelflow: dW f _ l.6MW f ~ δM ~ 1+.2M 2 δW f _ W f δW 'fr f „ W fr, m) Compute the total derivative of corrected fuelflow with respect to mach: dW fe = δx 6 (bl -\) 1 dc dM ~ 2a q dM n) Compute the total derivative of corrected fuelflow with respect to mach: dW J _ _ cW L cW f dW fc ~ dM ~ ~ a + ^ c ~ dM o) Compute the partial derivative of specific range with respect to mach: —R — where π J r. i-s crosswind δtA p) Compute the total derivative of specific range with respect to mach at the desired mach: dS R _ = $ R _ _^_ dW L = 8 R __ S R _ dW j _ dM ~ cM + W f dM ~ dM W f dM To determine maximum speed M ma _ at fixed altitude, described with reference to Figure 7B, the mach number M which drives T-D to T rc with fuelflow at the rated value dW t (W f = W fr ) is found. This is performed by "guessing" M and solving for W f and — — in a similar manner as done for the case of determining M wc . Performing this at two initial values of M allows the computation of M max to drive W f to equal W fr as shown in Figure 10. Maximum endurance speed, V me , Figure 7C, is first computed as: v meo = v ref +ιo and p 0 = standard sea level air density = .00237692 slugs/ft*^. C L — reference aircraft coefficient of lift-obtained from the specific aircraft data file. V me is then checked for being thrust limited as follows: where V TL , thrust-limited speed is taken from the mach number M which drives (T- E>)|„* to T rc as in the maximum speed case, the only difference being the smaller solution is chosen. In the special case where no solution exists, the M which maximizes (T- D)\ Vfi is used. The prediction of manual speed mode, or optimal A , or maximum altitude h pmxx for manual speed, Figure 7A, includes three submodes: a KCAS only, a Mach only, or both KCAS and Mach specified. Single dimension searches of S R and T- D are used to find h , and A altitude, respectively. In order to perform those searches, the — — and dh. -ϊ-~- mus, include .he ~-~ for ,he different speed submode, These can be derived dh P B dh P π v from standard expressions and using a standard lapse rate. The max altitude A search estimates the place where (T- D) = 0 as shown by Figure 11 A, whereas the opt altitude A , search finds where — - = 0 as shown in Figure 1 IB. dh In the higher dimension cases, where there are n variables, the following general method is employed for finding optimum altitude h popl and maximum altitude for long range cruise speed M wc , Figure 8B, or optimum altitude h popt for maximum speed M mΑ at rated fuelflow W fr , Figure 8C. Gradients must be computed at each initial guess; the number of initial guesses exceeding the number of dimensions n by at least 1. The gradients are then used to fit a "second order surface" to the real surface on which the summit or other point of interest on the best fit surface is found. As the one dimensional case depends on the first derivative being approximately linear, this higher dimension method, depends on similar restrictions on the gradient, such that: V/ * _4x +b which implies that the surface itself may be expressed as: f = -x T Ax +b T x + C = 0 for A = A 1 Having chosen initial guesses, x σ , a "Matrix of Gradients" is expressed as: VF Since each Vf 0 » Ax 0 +b v o therefore: Solving for C yields C = (M T M) ~X M r VF 0 // m > n + 1, or C = M ~λ VF 0 if m - n-v 1, A and b are then easily extracted from C. If the point of interest is the summit of the surface, this point can be found from: _4x +Z> * V/ = 0 then x, = -A ~l b is the best estimate of where V/ goes to zero. Using this method, maximum speed M m ... and optimum altitude h popl therefor, Figure 8C, are determined. Assuming that pressure altitude h p and true airspeed V τ are the only independent variables, such that weight W and temperature T D]SA are completely specified by h p and fuelflow W f is at the rated fuelflow value W fr , then W f is specified given h p and V τ . Defining a residual climb thrust T rc as before, T-D as a function of h p and V τ can be plotted in the form of a contour map as shown in Figure 12. The point on the contour T-D = T rc which maximizes V τ is the point at which maximum true airspeed occurs for any altitude subject to the constraint that the aircraft be able to initially maintain a rate of climb of 100 fpm. The coordinates of that point give the maximum airspeed M max and the altitude at which it will occur, the latter taken to be optimum altitude h popt for maximum speed M max . At any point on a given T-D contour, a tangent at that point is orthogonal to the gradient at that same point. Thus: and Tangent vectors = The slope of the contours at a given point can then be expressed as: The method for finding maximum speed M max and optimum altitude A , arises from dV τ noting that — - = 0 at maximum speed M max and optimum altitude A , . This dh p c(T- D) condition manifests itself on the contour map as = 0 which in conjunction ^P with the fact that T- D = T rc gives the information required to move toward a final solution in both axes, simultaneously. The following steps are then performed: 1) Choose 3 initial guess points: x ,x 0b and x 0c as follows: where h poa = crossover altitude for V mo I M mo . V τ = maximum speed at A (max speed, fixed alt case) ^ = ^ + °f x 0c = o + *[0θ ΔAV T -. where Δ ΔAA P^o = 4000' AV T =.05 mach expressed as true airspeed If M 0 , the mach number corresponding to V τ , is larger than M mo , M max is taken to be M mo and h popt to be the crossover altitude and the algorithm is complete. Otherwise, 2) Compute the gradient of the T-D surface at the first initial point: c τ dV If — — | A . < 0, V τ and h popt are found by solving the following simultaneous equations: p uπ p Vτ = Vτ , +V h (h -h where /is the well known function relating true airspeed to calibrated airspeed, altitude and temperature. -Η nax 1S tnen ta ken from V τ and the algorithm is complete. dV Otherwise, — -\ x ≥ 0, and the algorithm continues with step 3). dh. ° ~ 3) Assuming that (T-D) (x) can be approximately modeled as: = (T-D) 0 +b τ Ax + -Ax τ AAx Then the gradient is: VT(x) * b + -(A + A τ )Ax = b + AAx since A = A τ A and b can then be estimated from the values computed in Steps 3 and 4, and where VT(x 0t ) and VT(x 0c ) are completed in the same manner as V7(x σ ). C = M ~X VT and A and b are extracted from C. 4) The best estimate of the point x, = [A Λ V τ _ f (where V τ is maximum speed and h p _ is optimum altitude for maximum speed) is found from the constraints: Vr(χ. ) = (x. - x 0 ) + b = (T-D)(x l ) ~ - (T-D)(x + b τ (x ] -x + (x 1 -x τ A(x 1 - x θa ) - ~ 0 Two simultaneous equations, one linear and one quadratic, in A and V τ can then be written. The solution for which - . , < 0 is correct. ■iΨ τ Α -h"y\ 5) The point x, then becomes the new x 0^ . New guesses for x 0t and x 0c are chosen which have an order of magnitude less displacement from x 0 than the displacement used on the first iteration. Steps 1-5 are then iterated until the movement in mach is less than .003 and the movement in altitude is less than 100 ft. The long range cruise speed M wc and optimum altitude h pop! and maximum altitude h pπax for M mc , Figure 8B, is calculated in a similar manner as shown below: a) Maximum altitude h pmax is found first because if the optimum altitude h popt is limited by the maximum altitude h pma , then no further iterations are required to find optimum altitude h popt . b) A seed mach M seed and altitude h p is chosen as an initial guess: h p = 0.9 h ceιl where h ceιl is the ceiling altitude in ft. where S is wing area in ft 2 and δ is the atmospheric pressure ratio. c) Based on this initial guess, S Λ and (T-D) surfaces are computed utilizing general surface search techniques as described previously with respect to determination of maximum speed. The T-D surface is computed using rated fuelflow as an input, and this therefore represents the maximum excess thrust at a given speed M and altitude h p and is shown in Figure 14. d) Long range cruise speed M LRC is defined to be where S R is 99% of the peak S Λ at the high speed side of the peak as shown in Figure 13. e) The maximum altitude A for M LRC is the place where T-D ~ =0 intercepts the 99% S R maximum line as shown by Figure 14. f) The maximum specific rangeS Λ is given by the following two conditions: Equation I is solved to yield V τ as a function of A . Inserting V τ into Equation II yields an expression for S R as a function of h p . g) The intercept on the T-D surface, representing h p , Figure 14, satisfies the following two conditions: 0 = TD„ +B. TD IV By inserting the expression for S Λ as a function of h p into Equation III, the problem of estimating h p and V- is reduced to a standard two-equations-two-unknowns computational problem. h) Optimum altitude A^_, for M RC is then found. If at the maximum altitude h pmex so — — > 0, then h popt is set equal to A pmax . Otherwise, the A f is determined at the point where S R peaks with respect to altitude h p . This condition is written as: Once again, using Equation III and V, the problem becomes one of two-unknowns-two- equations. In many systems, it is difficult to say how accurate the predictions are after a fixed number of flights, which is dependent on the learning portion 12. This accuracy problem arises because the degree of accuracy depends on the kinds of flights that have been flown, i.e. an operator whose first ten flights have "visited" all corners of the flight envelope will have a fairly settled and accurate system, whereas an operator whose first ten flights were essentially duplicates of one another will have significant changes occur in the learning portion 12 when the operator finally operates in these corners of the flight envelope. The present invention, however, provides an observable indicator shown as figure of merit 24, Figure 1, which indicates accuracy of the current estimate of fuel remaining at the destination. The figure of merit, FOM, is determined utilizing the covariance matrix P along with Weight W τoc , altitude h p , speed M and fuelflow W f , all at top of climb, as follows: where H τ is comprised of the first seven terms of the normalized (T-D) H matrix evaluated with the aforementioned top-of-climb terms, P τ is the upper left 7x7 portion of the (T-D) covariance P, and t "E r I r E r is estimated time enroute Those skilled in the art will recognize that only preferred embodiments of the present invention have been disclosed herein, that other advantages may be found and realized, and that various modifications may be suggested by those versed in the art. It should be Uι. iβrstood that the embodiment shown herein may be altered and modified without departing from the true spirit and scope of the invention as defined in the accompanying claims.

Claims: CLAIMS 1. An estimating method for use in predicting performance characteristics of an aircraft, said method comprising the steps of: modeling the aircraft with at least one series expansion using aircraft specific data and input parameters which define the performance characteristics; and learning coefficients of said at least one series expansion over at least one flight so that said learned coefficients can be utilized for predicting performance characteristics of subsequent flights. 2. A method according to Claim 1, wherein said learning step includes the step of utilizing at least one filter to estimate said coefficients of said at least one series expansion. 3. A method according to Claim 1, wherein said modeling step includes the step of modeling a thrust -minus-drag relationship with a first series expansion. 4. A method according to Claim 3, wherein said first series expansion includes first terms representative of thrust and second terms representative of drag, said second terms including a compressible drag portion. 5. A method according to Claim 3, wherein said input parameters include altitude. 6. A method according to Claim 3, wherein said modeling step further includes the step of modeling fuelflow with a second series expansion. 7. A method according to Claim 6, wherein said learning step includes the step of utilizing a first filter for estimating coefficients of said first series expansion and a second filter for estimating coefficients of said second series expansion. 8. A method according to Claim 3, wherein said first series expansion is a Taylor series expansion. 9. A method according to Claim 6, wherein said first and second series expansions are Taylor series expansions. 10. A method according to Claim 7, wherein said first filter is an extended Kalman filter and said second filter is a Kalman filter. 11. A method according to Claim 3, wherein said learning step includes the step of estimating acceleration of the aircraft, said acceleration estimate is a function of estimated values of rate of change of airspeed and altitude. 12. A method according to Claim 11, wherein said terms of said first series expansion modeling said thrust and drag relationship are processed only when said acceleration estimate is within certain predetermined limits. 13. A method according to Claim 1, wherein said learning step includes the steps of estimating an accuracy of said coefficients of said at least one series expansion to provide information as to an accuracy of performance characteristics predicted in subsequent flights. 14. A method of predicting performance characteristics of an aircraft, said method comprising the steps of: modeling the aircraft with at least one mathematical model using aircraft specific data and input parameters which define the performance characteristics; learning coefficients of said mathematical model over at least one flight for a given flight plan; and predicting certain performance characteristics utilizing at least a first derivative taken with respect to at least one input parameter of a function derived from said at least one learned mathematical model given certain predetermined conditions and said learned coefficients. 15. A method according to Claim 14, wherein said modeling step includes the step of modeling a thrust-minus-drag relationship with a series expansion. 16. A method according to Claim 15, wherein said modeling step further includes modeling fuelflow with an additional series expansion. 17. A method according to Claim 16, wherein said learning step includes estimating said coefficients of said series expansion and said additional series expansion over at least one flight. 18. A method according to Claim 17, wherein said predicting step further includes predicting climb performance characteristics of the aircraft as a function of said series expansion and/or said additional series expansion. 19. A method according to Claim 18, wherein said predicting step further includes predicting descent performance characteristics of the aircraft as a function of said series expansion and/or said additional series expansion. 20. A method according to Claim 18, wherein said predicting steps include performing step integrations over said series expansion and/or said additional series expansion. 21. A method according to Claim 14, wherein said modeling step includes the steps of modeling a thrust-minus-drag relationship with a first mathematical model and modeling fuelflow with a second mathematical model. 22. A method according to Claim 21, wherein said learning step includes the steps of utilizing a first filter to estimate coefficients of said first mathematical model and utilizing a second filter to estimate coefficients of said second mathematical model. 23. A method according to Claim 22, wherein said modeling step includes the steps of modeling the thrust-minus-drag relationship with a first Taylor series expansion and modeling fuelflow with a second Taylor series expansion, and further wherein said learning step includes the steps of estimating coefficients of said first Taylor series expansion utilizing an extended Kalman filter and estimating coefficients of said second Taylor series expansion utilizing a Kalman filter. 24. A method according to Claim 14, wherein said certain performance characteristics include cruise performance characteristics. 25. A method according to Claim 24, wherein said predicting step includes the step of predicting long range cruise speed. 26. A method according to Claim 25, wherein said predicting step includes the step of predicting optimum altitude for said predicted long range cruise speed. 27. A method according to Claim 24, wherein said predicting step includes the step of predicting maximum speed. 28. A method according to Claim 27, wherein said predicting step includes the step of predicting optimum altitude for said predicted maximum speed. 29. A method according to Claim 24, wherein said predicting step includes the step of predicting optimum altitude at a manual speed. 30. A method according to Claim 24, wherein said predicting step includes the step of predicting maximum endurance speed. 31. A method according to Claim 14, wherein said learning step includes the step of estimating current acceleration of the aircraft. 32. A method according to Claim 14, wherein said learning step includes the step of utilizing altitude as an input parameter in defining the performance characteristics. 33. A method according to Claim 14, further including the step of predicting the accuracy of predictions of said certain performance characteristics. 34. A method according to Claim 14, wherein said learning step includes the step of utilizing previously learned data from a separate flight of other like aircraft to initialize coefficients of said mathematical model. 35. A method according to Claim 14, wherein said learning step includes the step of utilizing previously learned coefficients from a separate flight of other like aircraft for predicting performance characteristics of a current flight and given flight plan. 36. A method of predicting performance characteristics of an aircraft, said method comprising the steps of: modeling thrust minus drag of an aircraft with at least one mathematical model using aircraft specific data and input parameters which define the performance characteristics; and learning coefficients of said at least one mathematical model; and predicting the performance characteristics as a function of said learned coefficients representative of a combined and non-independent thrust -minus-drag relationship. 37. A method according to Claim 36, wherein said predicting step includes the step of predicting certain performance characteristics utilizing at least a first derivative of a function derived from said mathematical model with respect to at least one input parameter given certain predetermined conditions. 38. A method according to Claim 37, wherein said learning step includes the step of using learned coefficients from at least one separate flight of other like aircraft to initialize coefficients of said mathematical model. 39. A method according to Claim 36, wherein said method further includes estimating an accuracy of the performance characteristics predicted. 40. A method according to Claim 36, wherein said modeling step further includes the step of estimating modeling fuelflow with an additional mathematical model, said predicting step further includes the step of predicting performance characteristics as a function of said mathematical model and/or said additional mathematical model. 41. A method according to Claim 40, wherein said predicting step further includes predicting performance characteristics of the aircraft for a given flight plan utilizing a closed-form integration technique. 42. A method according to Claim 41, wherein said predicting step includes predicting time enroute, distance and/or total fuel burned for said given flight plan.

Inventors: RINGES ERIK A, SMITH JOSEPH A, TARBUTTON VICTOR F, RINGES, ERIK, A., SMITH, JOSEPH, A., TARBUTTON, VICTOR, F.
Applicants: HONEYWELL INC, HONEYWELL INC.
Application year: 1994
Main classification: G05D1
IPC codes: G05D1